A gas turbine engine typically includes a fan and a turbomachine. The turbomachine generally includes an inlet, one or more compressors, a combustor, and at least one turbine. The compressors compress air which is channeled to the combustor where it is mixed with fuel. The mixture is then ignited for generating hot combustion gases. The combustion gases are channeled to the turbine(s) which extracts energy from the combustion gases for powering the compressor(s), as well as for producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator.
In at least certain embodiments, the turbomachine and fan are at least partially surrounded by an outer nacelle. With such embodiments, the outer nacelle defines a bypass airflow passage with the turbomachine. Additionally, the turbomachine is supported relative to the outer nacelle by one or more outlet guide vanes/struts.
During operation of the gas turbine engine, various systems may generate a relatively large amount of heat. Thermal management systems of the gas turbine engine may collect heat from one or more of these systems to maintain a temperature of such systems within an acceptable operating range. The thermal management systems may reject such heat through one or more heat exchangers. In at least certain embodiments, at least one of these heat exchangers may be integrated into one or more components exposed to the bypass airflow passage, such as one or more of the struts extending between the turbomachine and the outer nacelle.
However, inclusion of one or more heat exchangers integrated into one or more components exposed to the bypass airflow passage may have an adverse effect on an airflow through the bypass airflow passage. Accordingly, a thermal management system capable of rejecting heat from various components of the gas turbine engine without adversely affecting the airflow through the bypass airflow passage would be useful.